Variable cycle compensation in a gas turbine engine

ABSTRACT

An aspect includes a variable cycle system of a gas turbine engine. The variable cycle system includes an actuation system, an electric component, and a controller. The actuation system is configured to adjust a variable cycle of turbomachinery of the gas turbine engine. The electric component is operable to provide a shaft power supply or a load corresponding respectively to an adjustment of the turbomachinery. The controller is operable to adjust an output of either or both of the actuation system and the electric component for separate control of thrust and cycle responses.

BACKGROUND

The subject matter disclosed herein generally relates to engine systemsand, more particularly, to a method and apparatus for variable cyclecompensation in a gas turbine engine.

Variable geometry in a gas turbine engine can provide a faster thrustresponse as one or more components of the gas turbine engine areadjusted in position or orientation as compared to only modifying a fuelflow rate to accelerate or decelerate a rate of engine spool rotationwithin the gas turbine engine. Engine thrust response changes aretypically slower to adjust relative to the rate at which variablegeometry of the engine can change.

BRIEF DESCRIPTION

According to one embodiment, a variable cycle system of a gas turbineengine includes an actuation system, an electric component, and acontroller. The actuation system is configured to adjust a variablecycle of turbomachinery of the gas turbine engine. The electriccomponent is operable to provide a shaft power supply or a loadcorresponding respectively to an adjustment of the turbomachinery. Thecontroller is operable to adjust an output of either or both of theactuation system and the electric component for separate control ofthrust and cycle responses.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the controlleris further operable to receive a control input, determine a plurality ofcurrent operating conditions of the gas turbine engine, calculate aplurality of commands to a plurality of power production and absorptionsubsystems for adjusting the variable cycle based on the currentoperation condition of the gas turbine engine using a plurality ofmodels of the subsystems that describe relationships between thecommands and respective impacts on engine power production andabsorption, and communicate the commands to the power production, powerabsorption, and one or more other actuation subsystems based on thecontrol input and the current operating conditions.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the electriccomponent is a motor-generator.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the electriccomponent absorbs power as an electric generator to produce electricalpower for an aircraft use or recharging of a battery system.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the actuationsystem includes a variable area turbine.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the electriccomponent adds power as an electric motor.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the actuationsystem includes one or more of: a variable area nozzle, a variable fanblade angle, an adaptive fan system, and a rotating fan inlet guidevane.

According to another embodiment, a propulsion system includes a gasturbine engine, an actuation system configured to adjust a variablecycle of turbomachinery of the gas turbine engine, and a variable cyclesystem compensation means operable to provide a shaft power supply or aload corresponding respectively to an adjustment of the turbomachineryfor separate control of thrust and cycle responses.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the variablecycle system compensation means performs control operations based on oneor more of: a thrust command, a throttle lever angle, a clearance, acompressor parameter, and a turbine parameter.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the variablecycle system compensation means includes a generator, and the actuationsystem includes a variable area turbine.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the variablecycle system compensation means includes an electric motor, and theactuation system includes one or more of: a variable area nozzle, avariable fan blade angle, an adaptive fan system, and a rotating faninlet guide vane.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the variablecycle system compensation means includes a motor-generator operable in agenerator mode to charge a battery system and in a motor mode to providesupplemental rotation force to the gas turbine engine based on electriccurrent from the battery system or an auxiliary power unit.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where the batterysystem is used during flight to power one or more electrical systems.

According to another embodiment, a method of variable cycle compensationin a gas turbine engine includes receiving a control input at acontroller of the gas turbine engine and determining a plurality ofcurrent operating conditions of the gas turbine engine. An output ofeither or both of an actuation system and an electric component isadjusted for separate control of thrust and cycle responses based on thecontrol input and the current operating conditions, where the actuationsystem is configured to adjust a variable cycle of turbomachinery of thegas turbine engine and the electric component is operable to provide ashaft power supply or a load corresponding respectively to an adjustmentof the turbomachinery.

In addition to one or more of the features described above or below, oras an alternative, further embodiments may include where determining theplurality of current operating conditions of the gas turbine engine isperformed using a control model and a plurality of actuator parameters,sensor values, and limits.

A technical effect of the apparatus, systems and methods is achieved byproviding variable cycle compensation in a gas turbine engine asdescribed herein.

BRIEF DESCRIPTION OF THE DRAWINGS

The following descriptions should not be considered limiting in any way.With reference to the accompanying drawings, like elements are numberedalike:

FIG. 1 is a partial cross-sectional illustration of a gas turbineengine, in accordance with an embodiment of the disclosure;

FIG. 2 is a schematic diagram of a variable cycle system, in accordancewith an embodiment of the disclosure;

FIG. 3 is a thrust plot of a snap acceleration, in accordance with anembodiment of the disclosure;

FIG. 4 is a speed plot of a snap acceleration, in accordance with anembodiment of the disclosure;

FIG. 5 is a fan pitch plot of a snap acceleration, in accordance with anembodiment of the disclosure;

FIG. 6 is a motor power plot of a snap acceleration, in accordance withan embodiment of the disclosure;

FIG. 7 is a thrust plot of a high frequency oscillation, in accordancewith an embodiment of the disclosure;

FIG. 8 is a speed plot of a high frequency oscillation, in accordancewith an embodiment of the disclosure;

FIG. 9 is a fan pitch plot of a high frequency oscillation, inaccordance with an embodiment of the disclosure;

FIG. 10 is a motor power plot of a high frequency oscillation, inaccordance with an embodiment of the disclosure;

FIG. 11 is a block diagram of a propulsion system, in accordance with anembodiment of the disclosure;

FIG. 12 is a flow chart illustrating a method, in accordance with anembodiment of the disclosure; and

FIG. 13 is a flow chart illustrating a method, in accordance with anembodiment of the disclosure.

DETAILED DESCRIPTION

A detailed description of one or more embodiments of the disclosedapparatus and method are presented herein by way of exemplification andnot limitation with reference to the Figures.

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct, while the compressorsection 24 drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a low pressure compressor 44 and a low pressureturbine 46. The inner shaft 40 is connected to the fan 42 through aspeed change mechanism, which in exemplary gas turbine engine 20 isillustrated as a geared architecture 48 to drive the fan 42 at a lowerspeed than the low speed spool 30. The high speed spool 32 includes anouter shaft 50 that interconnects a high pressure compressor 52 and highpressure turbine 54. A combustor 56 is arranged in exemplary gas turbine20 between the high pressure compressor 52 and the high pressure turbine54. An engine static structure 36 is arranged generally between the highpressure turbine 54 and the low pressure turbine 46. The engine staticstructure 36 further supports bearing systems 38 in the turbine section28. The inner shaft 40 and the outer shaft 50 are concentric and rotatevia bearing systems 38 about the engine central longitudinal axis Awhich is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion. It will be appreciated that each of the positions of the fansection 22, compressor section 24, combustor section 26, turbine section28, and fan drive gear system 48 may be varied. For example, gear system48 may be located aft of combustor section 26 or even aft of turbinesection 28, and fan section 22 may be positioned forward or aft of thelocation of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present disclosure isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,688 meters). The flight condition of 0.8 Mach and35,000 ft (10,688 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFC’)”—is the industry standard parameter of lbm of fuelbeing burned divided by lbf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]{circumflex over( )}0.5. The “Low corrected fan tip speed” as disclosed herein accordingto one non-limiting embodiment is less than about 1150 ft/second (350.5m/sec).

While the example of FIG. 1 illustrates one example of the gas turbineengine 20, it will be understood that any number of spools, inclusion oromission of the gear system 48, and/or other elements and subsystems arecontemplated. Further, rotor systems described herein can be used in avariety of applications and need not be limited to gas turbine enginesfor aircraft applications. For example, rotor systems can be included inpower generation systems, which may be ground-based as a fixed positionor mobile system, and other such applications.

FIG. 2 illustrates turbomachinery 202 that includes at least onecompressor section 204 and at least one turbine section 208 operablycoupled to a shaft 206 as part of an aircraft 200. The turbomachinery202 can be a spool of the gas turbine engine 20 of FIG. 1, such as thelow speed spool 30 or the high speed spool 32. For example, whenembodied as the low speed spool 30, the at least one compressor section204 can be equivalent to the low pressure compressor 44, the shaft 206can be equivalent to the inner shaft 40, and the at least one turbinesection 208 can be equivalent to the low pressure turbine 46 of FIG. 1.When embodied as the high speed spool 32, the at least one compressorsection 204 can be equivalent to the high pressure compressor 52, theshaft 206 can be equivalent to the outer shaft 50, and the at least oneturbine section 208 can be equivalent to the high pressure turbine 54 ofFIG. 1. The at least one compressor section 204 can include a fan, suchas an adaptive fan system or a fan that supports a variable fan bladeangle.

In the example of FIG. 2, a variable cycle system 210 is operablycoupled to the turbomachinery 202. The variable cycle system 210includes an electric component 212 operably coupled to the shaft 206. Inthe example of FIG. 2, a geared interface 230 operably couples theelectric component 212 to the shaft 206. The geared interface 230 caninclude, for instance, an auxiliary gear 233 coupled to an auxiliaryshaft 235 driven by the electric component 212. The geared interface 230can also include a rotor gear 237 coupled to the shaft 206. Theauxiliary gear 233 and the rotor gear 237 can each be beveled gears. Theauxiliary shaft 235 can be a tower shaft that enables the electriccomponent 212 to be separated at a greater distance from theturbomachinery 202 than direct coupling to the shaft 206 would provide.Further separation of the electric component 212 from the turbomachinery202 can improve accessibility to the electric component 212 forservicing and may reduce heating effects of the turbomachinery 202 onthe electric component 212 (e.g., due to fuel combustion). A disconnect240, such as a clutch, can be positioned between the electric component212 and a portion of the shaft 206 such that the electric component 212can be selectively engaged and disengaged to rotate with rotation of theshaft 206. In alternate embodiments, the electric component 212 isoperably coupled to the shaft 206 absent the geared interface 230 (e.g.,direct coupling).

The variable cycle system 210 also includes converter electronics 214operable to condition current to/from the electric component 212. Insome embodiments, the electric component 212 is a motor-generatorconfigurable in a generator mode to charge a battery system 250 and in amotor mode to provide supplemental rotation force to the turbomachinery202 of gas turbine engine 20 of FIG. 1. The electric component 212 caninclude conventional generator/motor components, such as a rotor andstator, including a plurality of windings and/or permanent magnets. Theconverter electronics 214 can also include conventional current controlelectronics, such as filters, switching components, rectifiers,inverters, voltage converters, and the like. The electric component 212can perform as a variable frequency generator in a generator mode due tospeed fluctuations of rotation of the shaft 206, which may be primarilydriven by the at least one turbine section 208. Alternatively, afrequency normalizing component can interface with the electriccomponent 212 to produce a constant frequency output (e.g., through theconverter electronics 214 or as a mechanical interface between theelectric component 212 and the shaft 206). In some embodiments, theelectric component 212 may be operable as a starter motor to partiallyor completely power rotation of the shaft 206 in a starting mode ofoperation (e.g., to start the gas turbine engine 20 of FIG. 1) and/orcan provide supplemental power to the shaft 206 during various flightphases of the aircraft 200. Other uses and functions for the electriccomponent 212 are contemplated.

The converter electronics 214 can control charging of the battery system250 responsive to a controller 216. The controller 216 can enable a flowof a charging current from the electric component 212 or a power input252 to charge the battery system 250 as regulated and conditionedthrough the converter electronics 214. The power input 252 can be anexternal input, such as power received through a plug interface whilethe aircraft 200 is on the ground at a ground-based power source, e.g.,at a gate or service location. In some embodiments, the converterelectronics 214 may receive electric current from an auxiliary powerinput 254 to provide a supplemental or alternative power source forcharging the battery system 250. For instance, the auxiliary power input254 may receive electric current from an auxiliary power unit (notdepicted) or another instance of the gas turbine engine 20 on theaircraft 200. The charge stored in the battery system 250 can provide anelectric current for a propulsion system use 256, which may includepowering one or more electric motors of the aircraft 200 during variousoperational states and/or providing power to the electric component 212when operating in a motor mode, for instance, to assist in drivingrotation of shaft 206. The propulsion system use 256 can be part of thegas turbine engine 20 that includes the turbomachinery 202 or anotheraircraft system, such as another instance of the gas turbine engine 20on the aircraft 200. The battery system 250 can be used on the ground orduring flight to power one or more electrical systems.

In embodiments, the controller 216 of the variable cycle system 210 canmonitor one or more rotor system sensors 218 while the turbomachinery202 is rotating. The rotor system sensors 218 can be any type orcombination of sensors operable to measure aspects of the motion of theturbomachinery 202. For example, the rotor system sensors 218 caninclude one or more accelerometers, speed sensors, torque sensors, andthe like. The rotor system sensors 218 can be existing sensors used forcontrolling the gas turbine engine 20. The controller 216 can control acharging of the battery system 250, for instance, by selecting thesource of electric current received through the converter electronics214. The controller 216 can also control operation of the electriccomponent 212. Data collected from the rotor system sensors 218 can beused to determine an operational status of a gas turbine engine 20 ofFIG. 2. Alternatively, the operational status of a gas turbine engine 20can be received as a signal or message from an alternate source, such asan engine system or aircraft communication bus. The controller 216 mayalso control other system aspects, such as controlling operation of thegas turbine engine 20 of FIG. 1. For example, the controller 216 can beintegrally formed or otherwise in communication with a full authoritydigital engine control (FADEC) of the gas turbine engine 20. The rotorsystem sensors 218 need not be directly coupled to the controller 216,as sensor data or sensor-derived data can be observed or determined byanother control (e.g., a FADEC) and provided to the controller 216. Inembodiments, the controller 216 can include a processing system 220, amemory system 222, and an input/output interface 224. The processingsystem 220 can include any type or combination of central processingunit (CPU), including one or more of: a microprocessor, a digital signalprocessor (DSP), a microcontroller, an application specific integratedcircuit (ASIC), a field programmable gate array (FPGA), or the like. Thememory system 222 can store data and instructions that are executed bythe processing system 220. In embodiments, the memory system 222 mayinclude random access memory (RAM), read only memory (ROM), or otherelectronic, optical, magnetic, or any other computer readable mediumonto which is stored data and algorithms in a non-transitory form. Theinput/output interface 224 is configured to collect sensor data from theone or more rotor system sensors 218 and interface with the converterelectronics 214 and/or other systems (not depicted).

The controller 216 may control an output of the electric component 212to compensate a shaft power supply and/or loading of the turbomachinery202 in coordination with an output of an actuation system 205. Theactuation system 205 can be located in a gas path of the gas turbineengine 20 FIG. 1. Although only a single instance of the actuationsystem 205, there can be multiple instances of the actuation system 205at various locations on the gas turbine engine 20 having differentfeatures that modify an engine geometry to impact thrust and/or otherparameters of the gas turbine engine 20. The actuation system 205 can bea means of adjusting power absorption, such as one or more of: avariable area nozzle, a variable fan blade angle, an adaptive fansystem, and a rotating fan inlet guide vane. To compensate powerabsorption, the electric component 212 can be implemented as an electricmotor or operated in a motor mode. Alternatively, the actuation system205 can be a means of adjusting power production, such as a variablearea turbine. The actuation system 205 can include any system operableto effect a variable cycle in the gas turbine engine 20, and may includecomponents not directly in the gas path, such as, an intercooled coolingair control system with an adjustable flow control. To compensate powerproduction, the electric component 212 can be implemented as an electricgenerator or operated in a generator mode. Thus, the variable cyclefeature of the actuation system 205 can be compensated by the electriccomponent 212. As one feature adjusts the ability of the gas turbineengine 20 to absorb power, the other provides the required power withouttime needed for spool speeds to change.

As a further example, the electric component 212 may be generallyreferred to as a variable cycle system compensation means operable tocompensate a power change induced by or in coordination with theactuation system 205, where the variable cycle system compensation meansis operable to respond at a second rate that is faster than a first rateof change initiated by the actuation system 205. The actuation system205 is a variable cycle component that may lag in responsiveness, forinstance, due to spool inertia and/or other factors. Further, althoughonly a single instance of the electric component 212 is depicted, therecan be multiple instances of the electric component 212 incorporated invariable cycle system 210, such as one or more dedicated instances of anelectric motor and an electric generator for one or more spools of thegas turbine engine 20.

FIGS. 3-6 illustrate a variable cycle compensation example for avariable fan paired with an electric motor with corresponding responsesduring a snap acceleration. For instance, a variable pitch fan canprovide a rapid thrust response without compensation, but power may bedrawn from spool inertia, which is counter-productive to increasingpower production of the turbine to compensate the increased powerabsorption of the variable pitch fan. Spool speed can go down, whichtakes time for the gas turbine engine 20 to catch up. By pairing anelectric motor as the electric component 212 with the variable pitch fanas the actuation system 205, the electric motor can compensate toincrease the ability of the variable pitch fan to translate shaft powerto thrust. This provides a more rapid thrust response without losingshaft speed. Thus, thrust response can be limited by actuator speedrather than spool speed.

FIG. 3 is a thrust plot 300 of a snap acceleration, in accordance withan embodiment. In thrust plot 300, a conventional engine with variablepitch fan response 302 is illustrated relative to an ideal response 304and a conventional engine acceleration response 306.

FIG. 4 is a low spool speed plot 400 of a snap acceleration, inaccordance with an embodiment. In low spool speed plot 400, aconventional engine with variable pitch fan response 402 is illustratedrelative to a variable pitch fan with electric motor response 404, and aconventional engine response 406.

FIG. 5 is a fan pitch plot 500 of a snap acceleration, in accordancewith an embodiment. In fan pitch plot 500, a conventional engine withvariable pitch fan response 502 is illustrated relative to a variablepitch fan with electric motor response 504 and a conventional engineresponse 506 that does not support a variable fan pitch.

FIG. 6 is a motor power plot 600 of a snap acceleration, in accordancewith an embodiment. In motor power plot 600, a conventional engine withvariable pitch fan response 602 is illustrated relative to a variablepitch fan with electric motor response 604 and a conventional engineresponse 606 that does not support a variable fan pitch. Collectively,the motor power depicted in the variable pitch fan with electric motorresponse 604 combines with the fan pitch in the variable pitch fan withelectric motor response 504 of FIG. 5 to enhance snap accelerationperformance as illustrated, for example, in FIGS. 3-6.

FIGS. 7-10 illustrate a variable cycle compensation example for avariable fan paired with an electric motor with corresponding responsesduring an auto-throttle or other high frequency oscillation event. Aconventional engine with variable pitch fan can see a rapid thrustresponse but dependence upon spool speed limits the ability to replicatea commanded high frequency thrust response. A pairing of an electricmotor as the electric component 212 with the variable pitch fan as theactuation system 205 can more accurately reproduce a commanded thrustprofile without affecting the cycle match. The electric motor can be amotor-generator and absorb excess power into energy storage and reducetotal energy use when the stored energy is harvested.

FIG. 7 is a thrust plot 700 of an auto-throttle, in accordance with anembodiment. In thrust plot 700, a conventional engine with variablepitch fan response 702 is illustrated relative to a variable pitch fanwith electric motor response 704. Notably, due to inertia effects, forexample, not only does the conventional engine with variable pitch fanresponse 702 fail to track the oscillation magnitude, but the delayedresponsiveness can result in a cumulative effect with each cycle ofoscillation. For instance, since the first peak value is not reached inthe conventional engine with variable pitch fan response 702, asubsequent cycle starts at a lower level which reduces the amplituderesponse for the conventional engine with variable pitch fan response702.

FIG. 8 is a low spool speed plot 800 of an auto-throttle, in accordancewith an embodiment. In low spool speed plot 800, a conventional enginewith variable pitch fan response 802 is illustrated relative to avariable pitch fan with electric motor response 804. As can be seen incomparison to FIG. 7, the imbalanced thrust response of the conventionalengine with variable pitch fan response 702 can result in an imbalancein the conventional engine with variable pitch fan response 802, withthe speed not held substantially constant. In contrast, the more rapidresponsiveness of the variable pitch fan with electric motor response704 can enable the variable pitch fan with electric motor response 804to remain substantially constant during the thrust oscillations of FIG.7.

FIG. 9 is a fan pitch plot 900 of an auto-throttle, in accordance withan embodiment, and FIG. 10 is a motor power plot 1000 of anauto-throttle, in accordance with an embodiment. In fan pitch plot 900,a variable pitch fan with electric motor response 904 is depicted. Inmotor power plot 1000, a variable pitch fan with electric motor response1004 is depicted. Collectively, the fan pitch in the variable pitch fanwith electric motor response 904 combines with the motor power depictedin the variable pitch fan with electric motor response 1004 to highfrequency oscillation performance as illustrated, for example, in FIGS.7-10.

Referring to FIG. 11, a block diagram of a propulsion system 1100 isdepicted in accordance with an embodiment. In the example of FIG. 11, anengine control 1104, such as controller 216 of FIG. 2, receives acontrol input 1102. The control input 1102 can be, for instance, one ormore of: a thrust command, a throttle lever angle, a clearance, acompressor parameter, and a turbine parameter. The control input 1102may be a pilot input or otherwise received on an aircraft bus. Theengine control 1104 may interface with multiple subsystems 1106, 1108,1110 to control an engine 1112, such as the gas turbine engine 20 ofFIG. 1. Power production subsystems 1106 can include either or both ofthe actuation system 205 and the electric component 212 of FIG. 2 whenoperating in a power production mode. Power absorption subsystems 1108can include either or both of the actuation system 205 and the electriccomponent 212 of FIG. 2 when operating in a power absorption mode.Primary actuation subsystems 1110 can include other non-cyclic actuationsystems, such as fuel system control, compressor vane control, and thelike. A control example for the engine control 1104 is depicted in FIG.12.

In FIG. 12, a method 1200 may be executed by the engine control 1104 ofFIG. 11 at each control time step in a loop. At block 1202, a controlinput 1102 of FIG. 11, such as a thrust command, can be communicated toand received at the engine control 1104.

At block 1204, the engine control 1104 can determine a plurality ofcurrent operating conditions of the engine 1112 of FIG. 11. Determiningthe plurality of current operating conditions of the engine 1112 can beperformed using a control model and a plurality of actuator parameters,sensor values, and limits.

At block 1206, a control problem solution can be generated. Examples ofcontrol system implementations can include constrained model-basedcontrol, multivariable system control, adaptive control, constraineddynamic inversion control, and other such techniques known in the art.Models can be incorporated for the power production subsystems 1106,power absorption subsystems 1108, and/or primary actuation subsystems1110 with associated limits and rate limits. For example, fan flow candepend on shaft speed and fan blade variable angle, which can be modeledbased on an actuator time response. Torque can be added or subtracted bythe electric component 212 of FIG. 2 from dynamical equations associatedwith spool dynamics, which can account for spool inertia. Models can beadded for electrical power generation and may include battery systemand/or auxiliary power unit effects. Rather than controlling to lowshaft speed, control objectives can be shifted to other parameters, suchas thrust, clearance, and/or other compressor or turbine parametersbeyond speed control. Compensation effects can be added and diminishedas balancing of power contributions or reductions shift between thepower production subsystems 1106 and the power absorption subsystems1108. Thus, the engine control 1104 can calculate a plurality ofcommands to a plurality of power production and absorption subsystems1106, 1108 for adjusting a variable cycle based on the current operationcondition of the gas turbine engine 20 using a plurality of models ofthe subsystems that describe relationships between the commands andrespective impacts on engine power production and absorption.

At block 1208, the engine control 1104 can communicate commands to powerproduction subsystems 1106, power absorption subsystems 1108, and/orprimary actuation subsystems 1110.

Referring now to FIG. 13 with continued reference to FIGS. 1-12, FIG. 13is a flow chart illustrating a method 1300 of variable cyclecompensation in a gas turbine engine, in accordance with an embodiment.The method 1300 may be performed, for example, by the variable cyclesystem 210 of FIG. 2 and propulsion system 1100 of FIG. 11. For purposesof explanation, the method 1300 is described primarily with respect tothe variable cycle system 210 of FIG. 2 and the propulsion system 1100of FIG. 11; however, it will be understood that the method 1300 can beperformed on other configurations (not depicted).

At block 1302, a control input 1102 can be received at a controller 216(e.g., engine control 1104) of a gas turbine engine 20 (e.g., engine1112). At block 1304, the controller 216 can determine a plurality ofcurrent operating conditions of the gas turbine engine 20. At block1306, the controller 216 can adjust an output of either or both of anactuation system 205 and an electric component 212 for separate controlof thrust and cycle responses based on the control input 1102 and thecurrent operating conditions. The actuation system 205 can be configuredto adjust a variable cycle of turbomachinery 202 of the gas turbineengine 20 (e.g., as the power production subsystems 1106 or the powerabsorption subsystems 1108). The electric component 212 can be operableto provide a shaft power supply or a load corresponding respectively toan adjustment of the turbomachinery 202.

While the above description has described the flow process of FIG. 13 ina particular order, it should be appreciated that unless otherwisespecifically required in the attached claims that the ordering of thesteps may be varied.

The term “about” is intended to include the degree of error associatedwith measurement of the particular quantity based upon the equipmentavailable at the time of filing the application.

The terminology used herein is for the purpose of describing particularembodiments only and is not intended to be limiting of the presentdisclosure. As used herein, the singular forms “a”, “an” and “the” areintended to include the plural forms as well, unless the context clearlyindicates otherwise. It will be further understood that the terms“comprises” and/or “comprising,” when used in this specification,specify the presence of stated features, integers, steps, operations,elements, and/or components, but do not preclude the presence oraddition of one or more other features, integers, steps, operations,element components, and/or groups thereof.

While the present disclosure has been described with reference to anexemplary embodiment or embodiments, it will be understood by thoseskilled in the art that various changes may be made and equivalents maybe substituted for elements thereof without departing from the scope ofthe present disclosure. In addition, many modifications may be made toadapt a particular situation or material to the teachings of the presentdisclosure without departing from the essential scope thereof.Therefore, it is intended that the present disclosure not be limited tothe particular embodiment disclosed as the best mode contemplated forcarrying out this present disclosure, but that the present disclosurewill include all embodiments falling within the scope of the claims.

What is claimed is:
 1. A variable cycle system of a gas turbine engine,the variable cycle system comprising: an actuation system configured toadjust a variable cycle of turbomachinery of the gas turbine engine; anelectric component operable to provide a shaft power supply or a loadcorresponding respectively to an adjustment of the turbomachinery; and acontroller operable to adjust an output of either or both of theactuation system and the electric component for separate control ofthrust and cycle responses.
 2. The variable cycle system of claim 1,wherein the controller is further operable to: receive a control input;determine a plurality of current operating conditions of the gas turbineengine; calculate a plurality of commands to a plurality of powerproduction and absorption subsystems for adjusting the variable cyclebased on the current operation condition of the gas turbine engine usinga plurality of models of the subsystems that describe relationshipsbetween the commands and respective impacts on engine power productionand absorption; and communicate the commands to the power production,power absorption, and one or more other actuation subsystems based onthe control input and the current operating conditions.
 3. The variablecycle system of claim 2, wherein the electric component is amotor-generator.
 4. The variable cycle system of claim 2, wherein theelectric component absorbs power as an electric generator to produceelectrical power for an aircraft use or recharging of a battery system.5. The variable cycle system of claim 2, wherein the actuation systemcomprises a variable area turbine.
 6. The variable cycle system of claim1, wherein the electric component adds power as an electric motor. 7.The variable cycle system of claim 1, wherein the actuation systemcomprises one or more of: a variable area nozzle, a variable fan bladeangle, an adaptive fan system, and a rotating fan inlet guide vane.
 8. Apropulsion system comprising: a gas turbine engine; an actuation systemconfigured to adjust a variable cycle of turbomachinery of the gasturbine engine; and a variable cycle system compensation means operableto provide a shaft power supply or a load corresponding respectively toan adjustment of the turbomachinery for separate control of thrust andcycle responses.
 9. The propulsion system of claim 8, wherein thevariable cycle system compensation means performs control operationsbased on one or more of: a thrust command, a throttle lever angle, aclearance, a compressor parameter, and a turbine parameter.
 10. Thepropulsion system of claim 8, wherein the variable cycle systemcompensation means comprises a generator, and the actuation systemcomprises a variable area turbine.
 11. The propulsion system of claim 8,wherein the variable cycle system compensation means comprises anelectric motor, and the actuation system comprises one or more of: avariable area nozzle, a variable fan blade angle, an adaptive fansystem, and a rotating fan inlet guide vane.
 12. The propulsion systemof claim 8, wherein the variable cycle system compensation meanscomprises a motor-generator operable in a generator mode to charge abattery system and in a motor mode to provide supplemental rotationforce to the gas turbine engine based on electric current from thebattery system or an auxiliary power unit.
 13. The propulsion system ofclaim 12, wherein the battery system is used during flight to power oneor more electrical systems.
 14. A method of variable cycle compensationin a gas turbine engine, the method comprising: receiving a controlinput at a controller of the gas turbine engine; determining a pluralityof current operating conditions of the gas turbine engine; and adjustingan output of either or both of an actuation system and an electriccomponent for separate control of thrust and cycle responses based onthe control input and the current operating conditions, wherein theactuation system is configured to adjust a variable cycle ofturbomachinery of the gas turbine engine and the electric component isoperable to provide a shaft power supply or a load correspondingrespectively to an adjustment of the turbomachinery.
 15. The method ofclaim 14, wherein determining the plurality of current operatingconditions of the gas turbine engine is performed using a control modeland a plurality of actuator parameters, sensor values, and limits. 16.The method of claim 14, wherein the electric component is amotor-generator.
 17. The method of claim 14, wherein the electriccomponent absorbs power as an electric generator to produce electricalpower for an aircraft use or charging of a battery system.
 18. Themethod of claim 14, wherein the actuation system comprises a variablearea turbine.
 19. The method of claim 14, wherein the electric componentadds power as an electric motor.
 20. The method of claim 14, wherein theactuation system comprises one or more of: a variable area nozzle, avariable fan blade angle, an adaptive fan system, and a rotating faninlet guide vane.